This invention relates to thermal barrier coatings for protecting internal components in a gas turbine engine from oxidation and corrosion during engine operation.
When a stream of incompletely burned atomized fuel droplets reaches the wall of the combustor in a gas turbine engine, a localized reducing atmosphere is created. This enables carbon deposits to form on the combustor wall. This condition usually occurs after the spray pattern of one or more fuel nozzles deteriorates, producing larger liquid fuel droplets. If the carbon deposits can bond to the combustor wall, large carbon nodules (several cubic centimeters in volume) can build up. Such localized reducing conditions can also cause carbon to form from fuel droplets prior to their collision with the wall. These small carbon particles can then bond upon impact with the wall, leading to carbon build-up. Periodic breaking off of pieces of these carbon deposits results in significant erosion damage to turbine airfoils, particularly to the first stage turbine blades, which impact with the carbon particles at speeds up to 2000 feet per second. Impact with turbine blades typically pulverizes the carbon nodules into much finer particles. Trailing edges of high-pressure turbine vanes and coatings on turbine shrouds are also damaged by grit blasting by high speed debris from pulverized carbon nodules.
Carbon bonding to the combustor wall is facilitated when the localized gaseous environment produced by the stream of impinging fuel droplets reduces carbide forming surface oxides. For example, for an uncoated superalloy combustor wall, reduction of chromium oxide permits chromium carbide to form, which bonds the carbon nodule to the combustor wall. Similarly, when a yttria stabilized zirconia thermal barrier coating is coated on the combustor wall, reduction of zirconium oxide permits zirconium carbide to form and bond the carbon nodule to the wall.
For the foregoing reasons, it would be desirable to provide some means for inhibiting the bonding of carbon nodules and carbon deposits to combustor walls in gas turbine engines.
More or less representative forms of thermal barrier coatings for use in gas turbine engines are described in U.S. Pat. No. 4,055,705 to Stephan Stecura and Curt Leibert, U.S. Pat. No. 4,248,940 to George Goward, Delton Gray and Richard Krutenat, U.S. Pat. No. 4,861,618 to Raymond Vine, Keith Sheffler and Charles Bevan, U.S. Pat. No. 5,073,433 to Thomas Taylor, and U.S. Pat. No. 5,514,482 to Thomas Strangman. These patents, however, make no mention of the carbon nodule problem and fail to suggest a solution to such problem.
In accordance with one feature of the invention, there is provided a carbon deposit inhibiting thermal barrier coating for an element (e.g., combustor wall) in a gas turbine engine. This coating comprises a layer of thermal barrier material formed on an exposed surface of a gas turbine engine element. This coating further comprises a layer of carbon deposit inhibiting material formed on top of the layer of thermal barrier material.
In accordance with another feature of the invention, there is provided an article for use in a gas turbine engine. Such article comprises a gas turbine engine element having a surface that will be exposed to burning engine gases and fuel droplets. Such article also includes a layer of thermal barrier material coated onto the engine element surface that will be exposed. This thermal barrier coating layer is typically composed of an insulative oxide layer and thin associated sublayers, such as an oxidation resistant bond coat that facilitates adhesion to the underlying surface. Such article further includes a layer of carbon deposit inhibiting material coated onto the outer surface of the thermal barrier material.
In accordance with a further feature of the invention, there is provided a method of forming a carbon deposit inhibiting thermal barrier coating on a gas turbine engine surface that will be exposed to the flow of burning engine gas and fuel droplets. Such method includes the step of depositing a layer of thermal barrier material onto the engine surface that will be exposed to the gas flow. Such method includes the further step of depositing a layer of carbon deposit inhibiting material onto the layer of thermal barrier material
For a better understanding of the present invention, together with other and further advantages and features thereof, reference is made to the following description taken in connection with the accompanying drawing, the scope of the invention being pointed out in the appended claims.